Combustion apparatus for gas turbine engines

ABSTRACT

A combustion chamber for a gas turbine engine which has staged combustion in two toroidal vortices of opposite hand arranged one upstream of the other. A burner delivers air/fuel mixture in a radial direction to support the vortices and the burner has a convergent outlet for the air/fuel mixture.

United States Patent Wood Sept. 23, 1975 [54] COMBUSTION APPARATUS FORGAS 3,430,443 3/1969 Richardson et al. 60/39.74 R TURBINE ENGINES3,724,207 4/1973 Johnson 60/39.74 R

[75] Inventor: Robert David Wood, Derby,

England [73] Assignee: Rolls-Royce (1971) Limited,

London, England [22] Filed: Sept. 4, 1973 [21] Appl. N0.: 394,395

[30] Foreign Application Priority Data Sept. 7, 1972 United Kingdom41504/72 [52] U.S. Cl 60/39.65; 60/39.7l; 60/39.74 R [51] Int. Cl. F02C7/22 [58] Field of Search 60/39.74 R, 39.65, 39.71

[56] References Cited UNITED STATES PATENTS 2,999,359 9/1961 Murray60/39.74 R

Primary ExaminerC. J. Husar Assistant ExaminerRobert E. GarrettAttorney, Agent, or Firm-Cushman, Darby & Cushman [57] ABSTRACT Acombustion chamber for a gas turbine engine which has staged combustionin two toroidal vortices of opposite hand arranged one upstream of theother. A burner delivers air/fuel mixture in a radial direction tosupport the vortices and the burner has a convergent outlet for theair/fuel mixture.

2 Claims, 6 Drawing Figures US Patent Sept. 23,1975 Sheet 1 of53,906,718

FIG!

US Patent Sept. 23,1975 Sheet 3 of5 FIGL US Patent Sept. 23,1975 Sheet 50f 5 3,906,718

COMBUSTION APPARATUS on GAS TURBINE ENGINES This invention relates tocombustion apparatus for gas turbine engines in which the fuel suppliedto a combustion chamber of the engine is dischargedfor combustion withinthe chamber.

It has been proposed to provide staged combustion for combustionchambers in order to improve combustion efficiency particularly inaircraft gas turbine engines at ground idling speed and to reduce smokeemission particularly at take-off. Such staged combustion has beenachieved by delivering the fuel into the chamber already mixed with aproportion of air and causing this air/fuel mixture to be discharged soas to create a substantial toroidal vortex at the upstream end of thecombustion chamber. At the same time, further secondary air has beenintroduced into the chamber downstream of the first toroidal vortex sothat the secondary air causes a second substantially toroidal vortex tobe generated which is of opposite hand to the first vortex.

During engine starting, the air/fuel mixture in the first toroidalvortex is of such strength as to be readily ignited by ignition meansmounted in the region of the vortex. As the fuel flow to the combustionchamber increases upon throttle opening, the first toroidal vortexbecomes over-rich in fuel and the majority of combustion is consequentlytransferred to the second toroidal vortex. When the fuel flow to thecombustion chamber decreases as the throttle is closed, combustion istransferrcd back to the first toroidal vortex as a result of theweakening of the air/fuel ratio in the vortices.

It has further been proposed to mix the fuel and air by discharging thefuel into a burner tube where it is mixed with air flowing .in thegeneral downstream direction.

The air/fuel mixture impinges upon a deflection member positioned at thedownstream end of the tube which deflects the mixture transversely so asto generate the said two vortices.

It has also been proposed to divide the tube of such a combustionapparatus into a number of passages extending axially of the engine soas to prevent the air/fuel mixture from swirling and to cause it to passdown the tube with substantially axial flow.

There are problems with staged combustion in that the flame tends tostabilise against projections or within recesses in the combustionchamber and this causes local burning of the chamber wall. This occursparticularly when fuel flow is being reduced and'combustion istransferring to the first upstream vortex. There are also problems infinding a design of burner which will eject the air/fuel mixture with asubstantially radial flow so as to support the two vortices at allcombustion rates.

The present invention is intended to overcome the problems associatedwith staged combustion so as to give complete mixing of the fuel and airin the tube and so as to cause the mixture to be discharged in adirection substantially radially of the axis of the burner whereby ithas been found that the double vortex combustion arrangement hasconsiderable stability over the whole range of operating conditions inthe engine.

In addition it has been found that the burner of the present inventiondoes not provide any projection into the combustion chamber upon which aflame can stabilise so that the burner is not destroyed rapidly and hasa useful life.

According to the invention there is provided combustion apparatuscomprising a combustion chamber, first and second'air inlet means fordelivering air into said chamber, said first air inlet meanscommunicating with a passage and means for supplying fuel to thepassage, a deflecting member disposed for deflecting the air/fuelmixture at the downstream end of the passage in a directionsubstantially radially of said passage whereby to promote asubstantially toroidal vortex circulating first radially outwards fromsaid passage, then towards the upstream end wall of the combustionchamber and then downstream adjacent and substantially parallel to thepassage, said secondary air inlet means being positioned to promote asecondary toroidal vortex immediately downstream of said first vortexand having a hand opposite to that of the first vortex, characterised inthat said passage is defined by a portion of a cylindrical member andsaid member forms with said deflecting member a convergent outlet fordirecting the air/fuel mixture in the direction substantially radiallyof said passage.

The invention is illustrated, merely by way of example, in theaccompanying drawings, in which:

FIG. 1 is a partly sectioned side elevation of a gas turbine engineprovided with combustion apparatus in accordance with the presentinvention,

FIG. 2 is a partial longitudinal section of the combustion apparatus ofthe engine of FIG. 1, to an enlarged scale,

FIG. 3 is a section taken on line III-III of FIG. 2,

FIG. 4 is an enlarged detail of part of FIG. 2,

FIG. 5 is a part longitudinal section of the combustion apparatus ofanother gas turbine engine which embodies an alternative form of theinvention, and,

FIG. 6 is a section taken along the line Vl-VI of FIG. 5.

With reference to FIG. 1, a gas turbine engine generally indicated at 10comprises, in axial flow series, an air intake 11, compressor means 12,combustion apparatus l3, turbine means 14 and an exhaust nozzle 15.Combustion apparatus 13 comprises a plurality of separate, substantiallycylindrical combustion chambers, one of which can be seen at 16,circumferentially mounted around the axis of the engine 10 in the mannerwhich is generally described as a cannular array.

The construction of the upstream portion of combustion chamber 16 can beseen more easily in FIG. 2. The combustion chamber 16, which isinterposed between the radially inner and outer walls 17 and 18 of anair casing of the engine 10, is positioned so as to receive aircompressed by the compressor means 12 via the duct 19. A combustor head20 of the combustion chamber 16 is attached to the remainder of thechamber via a cranked cross-section ring 21. The centre of the combustorhead 20 is apertured so as to provide a location for a burner 22 whichis secured to the head 20.

The burner 22 provides an air/fuel mixture which issues radially of theend of the burner in the direction of arrow 99 to create a firsttoroidal vortex 100. This vortex 100 is ignited and burning air/fuelmixture is entrained in a second toroidal vortex 200 which is generatedin part by the flow from burner 22 and in part by dilution air enteringthe chamber. The burning mixture passes from the second vortex in agenerally downstream direction to the downstream exit of the combustionchamber.

The burner 22 includes an outer shroud member 23 within which is mounted.a burner tube 24. The upstream end of the shroud member 23 defines aprimary convergent air inlet and a plurality of air holes 26. The burnertube 24 is mounted within the shroud member such that air from theprimary air inlet 25 passes into tube 24 whilst air from the holes 26passes along the external surface of tube 24.

The burner tube 24 consists of a cylindrical member 27 which encloses ahollow axially extending core portion 28 having five radially extendingvanes 29 formed along its length to define five axial passages 30, whichcan be seen more easily in FIG. 3. Air entering inlet 25 is caused todivide into five parallel flows and the ar rangement prevents any swirlcomponent being imparted to the air passing through the burner tube 24.The radially inner portions of the upstream ends of the vanes 29 arechamfered at 29a so as to accommodate a nozzle head 31 of a fuelinjector 32. The nozzle head 31 is provided with five fuel jets 33, eachfuel jet 33 being associated with a corresponding axially extendingpassage 30. The fuel jets 33 are positioned so as to direct jets of fuelinto the axial passages 30 in generallyradial directions which have aslight downstream component i.e., approximately normal to the flow ofair through the axial passages 30. A splash ring 34 is attached to theupstream ends of the vanes 29 such that it is co-axial with the axis ofthe core portion 28 and its perimeter is radially approximately mid-waybetween the axis of the core portion 28 and the periphery of the burnertube 24. The splash ring 34 is axially positioned such that the jets offuel discharged from the nozzle head 31 will impinge on it and will tendto become atomised and entrained in the air flowing along tube 24.

Any wakes in or maldistribution of that portion of air which passes intothe primary air inlet 25 are suppressed by the convergent configurationof the inlet. The primary air then passes over the nozzle head 31 wherethe fuel which is issued from jets 33 across the air flow is partiallyatomised. The portion of each of the fuel jets which is not immediatelyatomised impinges on the splash ring 34, thereby resulting in more fuelatomisation. Any remaining fuel which has not been atomised leaves thesplash ring 34 in the form of a fuel film. Atomisation of this fuel filmis effected by the interaction of the airflows which pass inside andoutside the splash ring 34.

The downstream end of the burner 22 is shown en larged in FIG. 4. Themixture passing down member 27 is confined in the five passages 30 in asubstantially annular channel of uniform radial extent. At thedownstream end of the member 27, the member is formed with an inwardlydirected lip 35 which causes any fuel droplets adhering to the insidewall of the tube 27 to be given aradially inwards motion so as to beentrained in the air flowing along the tube. The burner 22 includes adeflecting member 37 which has an upstream facing surface 40 which,together with the inside surfaccof member 27 forms a smooth convergentoutlet for the air/fuel mixture passing out of member 27. At theinwardly directed lip 35 the radial extent of this outlet is reduced andthe flow is turned radially outwards and passes to an annular convergentportion 36 before being further turned to emerge through opening 41. Theconvergent portion 36 includes successively reducing cross-sectionsdimensioned A and B. These are followed by a turning of the flow and asmooth convergence to the annular opening 41 from which the mixtureis'emitted'in a direction substantially radially with respect to thefore and aft direction of the engine and burner. Theouter portions ofmember 37 adjacent opening 41 are so shaped as to give the mixtureimpinging on them a substantially upstream and radially outwardcomponent of flow, which together with the downstream motion of theremaining flow, causes the flow out of opening 41 to be substantiallyradial.

The opening 41 is-defined by a lip 43 and the outermost portion 37a ofthe member 37 and is an annular opening having its plane lyingsubstantially parallel to the longitudinal axis of the engine. Thedisposition of opening 41 assists in ensuring that the flow out of theburner emanates radially.

The deflecting member 37 has a cylindrical portion 38 which locates inthe hollow centre of the core portion 28. The cylindrical portion 38 isthreaded at the end remote from the deflecting member 37 in order thatit may be retained within the core portion 28 by means of a nut 39. Theair which enters the holes 26 and passes through the annular ductdefined by the shroud member 23 and the burner tube 24 assists in thecooling of the burner tube 24. The shroud air issuing from this annularduct passes over the radially outer surface of the cylindrical member 27and impinges on the air/fuel mixture discharged from the annular opening41, thereby urging the mixture in a direction substantially radially ofthe central axis of the cylindrical member 27 and hence into thesubstantially toroidal vortex 100. The vortex circulates towards thecombustor head 20 of the combustion chamber 16 and then back along theradially outer surface of the shroud member 23 towards the opening 41,the shroud air passing over the radially outer surface of thecylindrical member serving to reinforce the vortex. The radiallyextending lip 43 is provided on the downstream edge of the cylindricalmember 27 in order to suppress any incipient turbulence at theconfluence of the shroud air and air/fuel mixture flows. The diameter ofthe lip 43 and the deflecting plate 37 are arranged to be substantiallyequal in order to avoid any tendency of the sharp edge of the deflectingmember 37 to burn. The combustion chamber 16 has a series of coolingstrips which are fed air through holes 42which produce film cooling onthe inside of the chamber to prevent heat degradation of the chamberwall.

Part of the portion of air which passes around the external surface ofthe combustion chamber 16 enters the chamber via a plurality ofsecondary air inlets 44 provided downstream of the deflecting member 37.As the axes of the secondary air inlets 44 are approximately normal tothe air flow passing thereover, that air which enters the secondary airinlets 44 does so obliquely and helps to support the secondsubstantially toroidal vortex 200. The vortex 200 thus formed liesimmediately downstream of the first formcd'vortex 100 and is ofoppositehand.

Thus two substantially toroidal vortices are formed in the combustionchamber 16 thereby permitting staged combustion. as hereinbeforedescribed, to take place.

By ensuring that the air which passes through the burner tube 24 does sowithout having a swirl component imparted to it, then much higher tubevelocities can be achievedthereby resulting in a very high relativeair/fuel velocity and consequently more efficient fuel atomisation.Similarly as the greatest air pressure drop occurs across the annularopening 41, then the exhaust velocity of the air/fuel mixture from theopening 41 is relatively high thereby increasing fuel atomisation andmixing efficiency even further and also preventing flame stabilization,and consequently burning of the burner 22 in the region of the annularopening 41. As more efficient fuel atomisation and mixing results in acorresponding increase in combustion efficiency, then combustionchambers in accordance with the present invention may have increasedcombustion efficiency and decreased smoke emission.

Referring now to FIGS. 5 and 6, there is shown an annular combustionchamber of a gas turbine engine, that is to say the combustion chamberis disposed in a single annulus concentric with the fore and aft centreline of the engine. A series of burners 50 is disposed around the innerperiphery of the combustion chamber 51. Each burner has a convergentinlet 52 for receiving air delivered to the combustion chamber alongpassage 53 and through a nose aperture 54. Each burner has apart-cylindrical tube 55 which is identical in respect of its downstreamproportions to the tube 27 of FIGS. 2 and 3. The burner includes adeflecting member 56 which deflects the air/fuel mixture passing downtube 55 in an outwardly radial direction, the shape of the inner surfaceof member 56 being similar to the surface 40 of the burner of FIGS. 2and 3. Thus the burner 55 provides a convergent outlet for the air/fuelmixture.

Fuel is delivered to the burner along a pipe 57 and discharged through afuel jet 58.

The burner 55 is partially enclosed by a shroud 60 and air passesbetween the shroud 60 and burner 55 to cool the outer surface of theburner and to assist the rotation of the toroidal vortex 100.

Adjacent the downstream end of the member 56 the inner wall of thecombustion chamber has a lip 61 which defines an aperture to allow aflow of air to cool member 56 and additionally to support the motion ofthe two vortices.

The combustion chamber includes secondary air openings 62 in its outerwall disposed to provide air which in part supports vortex 200. Inaddition the chamber includes cooling strips 63 which deliver a film ofcooling air to the inside wall of a chamber in a generally upstreamdirection over the area where the upstream flow of vortex 100 wouldimpinge on the chamber wall. The chamber also includes cooling strips 64which provide downstream flowing cooling films over the area where theflow of vortex 200 would impinge on the wall.

The particular arrangement of burners 50 to suit an annular combustionchamber is illustrated in FIG. 6. Each burner provides an outward radialflow and although all these flows are not precisely radial with respectto the axis of the engine the flows from adjacent burners merge to makethe overall flow substantially radial. The part cylindrical tube 55 maybe divided axially into a plurality of passages to prevent swirl and inthis arrangement each passage has its own fuel jet 58.

In an alternative arrangement (not illustrated) of the invention appliedto an annular combustion chamber the burners 55 are disposed around theouter periphery of the chamber. This avoids the use of the long fuelpipes 57 to each burner 55.

What I claim is:

1. Combustion apparatus comprising:

a combustion chamber having an upstream end wall and a downstream wall,openings in said chamber walls defining first and second air inlet meansfor delivering air into said chamber, a member extending axiallydownstream of said upstream end wall defining at least in part a passagewhich is in com munication with said first air inlet means, means forsupplying fuel to the passage, a deflecting member disposed fordeflecting the air/fuel mixture at the downstream end of the passage ina direction substantially radially of said passage whereby to promote asubstantially toroidal vortex circulating first radially outward fromsaid passage, then upstream toward the upstream end wall of thecombustion chamber and then downstream adjacent and substantiallyparallel to the passage, said secondary air inlet means being positionedto promote a secondary toroidal vortex immediately downstream of saidfirst vortex and having a hand opposite to that of said first vortex,said member defining said passage being a cylindrical member which formswith said deflecting member an annular outlet, and a plurality ofaxially extending straight dividers which divide the passage into aplurality of axially extending straight sectors and which prevent swirlin the air/fuel mixture passing through the passage sectors wherein saidoutlet converges radially around its circumference and terminates in anannular discharge opening through which the air/fuel mixture isdischarged into the combustion chamber.

2. Combustion apparatus comprising a combustion chamber having anupstream end wall and a downstream wall, openings in said chamber wallsdefining first and second air inlet means for delivering air into saidchamber, a member extending axially downstream of said upstream end walland defining at least in part a passage which is in communication withsaid first air inlet means, means for supplying a fuel to said passage,a deflecting member disposed for deflecting the air/fuel mixture at thedownstream end of the passage in a direction substantially radially ofsaid passage to thereby promote a substantially toroidal vortexcirculating first radially outward from said passage, then upstreamtoward the upstream end wall of the combustion chamber and thendownstream adjacent and substantially parallel to said passage, saidsecondary air inlet means being positioned to promote a secondarytoroidal vortex immediately downstream of said first vortex and having ahand opposite to that of said first vortex, said member defining saidpassage including a plurality of axially extending straight dividerswhich divide the passage into a plurality of axially extending straightsectors and which prevent swirl in the air/fuel mixture passing alongthe passage sectors, said member defining said passage forming with saiddeflecting member an outlet wherein said outlet converges radially aboutits circumference and terminates in a circumferential discharge openingthrough which the air/fuel mixture is discharged into the combustionchamber.

1. Combustion apparatus comprising: a combustion chamber having anupstream end wall and a downstream wall, openings in said chamber wallsdefining first and second air inlet means for delivering air into saidchamber, a member extending axially downstream of said upstream end walldefining at least in part a passage which is in communication with saidfirst air inlet means, means for supplying fuel to the passage, adeflecting member disposed for deflecting the air/fuel mixture at thedownstream end of the passage in a direction substantially radially ofsaid passage whereby to promote a substantially toroidal vortexcirculating first radially outward from said passage, then upstreamtoward the upstream end wall of the combustion chamber and thendownstream adjacent and substantially parallel to the passage, saidsecondary air inlet means being positioned to promote a secondarytoroidal vortex immediately downstream of said first vortex and having ahand opposite to that of said first vortex, said member defining saidpassage being a cylindrical member which forms with said deflectingmember an annular outlet, and a plurality of axially extending straightdividers which divide the passage into a plurality of axially extendingstraight sectors and which prevent swirl in the air/fuel mixture passingthrough the passage sectors wherein said outlet converges radiallyaround its circumference and terminates in an annular discharge openingthrough which the air/fuel mixture is discharged into the combustionchamber.
 2. Combustion apparatus comprising a combustion chamber havingan upstream end wall and a downstream wall, openings in said chamberwalls defining first and second air inlet means for delivering air intosaid chamber, a member extending axially downstream of said upstream endwall and defining at least in part a passage which is in communicationwith said first air inlet means, means for supplying a fuel to saidpassage, a deflecting member disposed for deflecting the air/fuelmixture at the downstream end of the passage in a directionsubstantially radially of said passage to thereby promote asubstantially toroidal vortex circulating first radially outward fromsaid passage, then upstream toward the upstream end wall of thecombustion chamber and then downstream adjacent and substantiallyparallel to said passage, said secondary air inlet means beingpositioned to promote a secondary toroidal vortex immediately downstreamof said first vortex and having a hand opposite to that of said firstvortex, said member defining said passage including a plurality ofaxially extending straight dividers which divide the passage into aplurality of axially extending straight sectors and which prevent swirlin the air/fuel mixture passing along the passage sectors, said memberdefining said passage forming with said deflecting member an outletwherein said outlet converges radially about its circumference andterminates in a circumferential discharge opening through which theair/fuel mixture is discharged into the combustion chamber.